The invention relates to an air inlet, particularly a two-dimensional air inlet set at an angle on one side for gas turbine jet propulsion plants for driving of airplanes, having an air inlet channel curved in space, especially doubly curved, and leading to the compressor of the propulsion plant, and further having, especially one, inlet cross-section for the supersonic operation controllable by means of adjustable ramps arranged overhead which are in the raised position during subsonic operation and which are tilted down during supersonic operation, so that a variable, convergent-divergent air inlet geometry is produced in the air inlet.
The air inlet of an aircraft has the purpose to convert as large as possible a proportion of the kinetic energy of the onflowing air into pressure energy while reducing its speed. This energy recovery may be substantial particularly at higher flight mach numbers. In order to optimize the energy recovery, the collected and compressed air must be supplied at low loss and in a homogeneous condition as well as in a properly dosed quantity to the propulsion plant. Flow losses result primarily due to air friction and compression shocks.
Particular attention must thereby also be paid to the resistances of the external flow which must be kept as low as possible.
During starting and at low flight speeds the air is supplied to the air propulsion plant with a large volume due to the low pressure at the inlet end. This means that the mechanically narrowest flow cross-section of the air inlet must be dimensioned to be large. Contrary thereto, at high supersonic flights the air volume at the inlet end is extremely low due to the produced high pressure. This means that the mechanically narrowest flow cross-sectional area must be dimensioned to be as small as possible in order to maintain the desired position of the compression shocks in the entrance zone of the air inlet.
Additionally, it is today required of high performance aircraft having a supersonic capability to be suitable for air combat in the subsonic range. For this purpose it is necessary to operate at high aircraft angles of attack which cause a slanted onflow. Such slanted onflow causes a flow separation or stalling in front at the lower lip of the air inlet which cause separations near the bottom which extend far into the air inlet channel or into the subsonic zone of the diffuser.
This adverse flow condition is particularly critical in its effect where the propulsion plants are installed in the aircraft fuselage and where the supersonic air inlets are arranged laterally of the aircraft fuselage. In this instance the doubly bent air supply channels from the air inlet to the respective propulsion plant aggrevate the adverse flow condition due to the flow phenomena to be described in the following.
The undisturbed flow which is present above the above mentioned separation flow near the bottom which is subject to large losses, runs, due to the inertia of its mass, with high speed against the inner surface of the channel wall which is bent toward the aircraft fuselage and forms an outer channel wall. Stated differently this undisturbed flow runs against the inner surface of the outer channel wall. Thus, a reduced pressure zone is formed in the area of the opposite inner flow wall and this reduced pressure zone calls for being filled aerodynamically. This is done by the separation flow near the bottom which contains less kinetic energy. The separation flow, due to its lateral flow off movement within the air inlet toward the inner flow wall on which it slides up, excites a spin flow. The energy of this spin flow is further amplified due to the fact that the flow present on the inside of the outer flow wall flows into the zone near the bottom which now is becoming empty, whereby this undisturbed flow displaces the flow which originally was separated near the bottom, further against the inner channel wall. A further intensification of this spin phenomenon may be caused by the following counter curvature of the air inlet channel. This is so because the undisturbed flow which up to this point was hugging the inner surface of the flow outer wall, is carried to the other wall of the air inlet channel downstream of the turning point in the curvature. Thus, the other wall now becomes the flow outer wall, whereby the original separated flow near the bottom is displaced upwardly toward the inner flow wall of the counter bend.
Today, gas turbine jet propulsion plants for high performance aircraft are altogether equipped with a multi-stage axial compressor. On the one hand, the advantages of an axial compressor are uncontested. However, on the other hand, an axial compressor has only a narrow stable working range and it is rather sensitive in its reaction to air inhomogenities. This weakness of the axial compressor may be counteracted by dividing the same into a plurality of compressor groups having different r.p.m.s. Further, this weakness may be counteracted by adjusting the guide vanes and by bleeding of compressor air between individual compressor stages. By these measures it is possible to avoid to a large extent the dreaded compressor pumping which occurs due to strong irregularities in the flow. When compressor pumping occurs, the air flow is interrupted between the individual compressor stages which in turn may lead to a substantial power reduction in the propulsion plant even to a total collapse of the propulsion process.
It follows from what has been said above that the propulsion plant of an aircraft comprising the air inlet and the gas turbine engine produces a thrust with good efficiency in all power ranges only if the air inlet and the engine cooperate aerodynamically in a stable manner. As mentioned, this may not be accomplished at all times under certain conceptual circumstances of the air inlet and under extreme operating conditions as in the present instance when flying with large aircraft angles of attack.